This sub-module complies with Part-66 11.1.1 / CAT B1.1/1 ...€¦ · Aviation Module 11 Aerodynamics, Structures & Systems Module 11.1.1 Aeroplane Aerodynamics & Flight Controls/MAR06

Aviation

Module 11 Aerodynamics, Structures & Systems

Module 11.1.1 Aeroplane Aerodynamics & Flight Controls/MAR06 Edition 1/Rev 0 TEC Training Manual TM11.1.1-2 page 1

This sub-module complies with Part-66 11.1.1 / CAT B1.1/1.2 TABLE OF CONTENT: Subject: Page: 11.1.1 Aeroplane Aerodynamics & Flight Controls ............................... 4 Aircraft Control ...................................................................................... 6

Primary Flight Controls................................................................... 6 Secondary Flight Controls.............................................................. 6

Control Handles and Levers ................................................................. 8 Pitch and Roll Controls .................................................................. 8 Yaw Control ................................................................................... 8 Trim Controls ................................................................................. 8 Horizontal (Movable/Trimable) Stabilizer Control........................... 8 Flaps and Slats Control.................................................................. 8 Spoiler and Speed Brake Control................................................... 8

The Function of Flight Control Surfaces ............................................. 10 Aerodynamic Principles................................................................ 10 Maneuverability ............................................................................ 10 Mechanical Stops......................................................................... 10 Hinge Moments............................................................................ 12 Direct Mechanical Control ............................................................ 12 Aerodynamic Balancing ............................................................... 12 Tab Balancing .............................................................................. 12

Powered Flight Controls...................................................................... 16 Power Boosted Control ................................................................ 16 Fully Powered Control.................................................................. 16 Artificial Feel & Centering ............................................................ 16 Fly-By-Wire Principle.................................................................... 16

Longitudinal (Pitch) Control................................................................. 18 Elevators ...................................................................................... 18 All Moving Tail Planes (Stabilizers).............................................. 18 All-Flying Tails.............................................................................. 20

Directional (Yaw) Control .................................................................... 22 Turn Coordination ........................................................................ 22 Yaw Damping............................................................................... 22

Engine Failure Compensation...................................................... 22

Lateral (Roll) Control........................................................................... 24 Aileron Roll Control ...................................................................... 24 Roll Spoilers................................................................................. 24 Aileron Reversal........................................................................... 24

Adverse yaw ....................................................................................... 26 Lift Inducing Devices........................................................................... 28

Augmentation of CLMAX ................................................................. 28 Trailing Edge Flaps ...................................................................... 28 Leading Edge Devices ................................................................. 32 Slots and Slats ............................................................................. 32 Leading Edge Flaps (Droop Nose) .............................................. 34 Krueger Flaps & Variable Camber LE-Flaps................................ 34 Combining Slats & LE-Flaps ........................................................ 34

Drag Inducing Devices........................................................................ 36 Spoilers ........................................................................................ 36 Speed-brakes............................................................................... 36 Drag Chutes................................................................................. 36

Combined Flight Controls ................................................................... 38 Flight Control Protections.................................................................... 40

Rudder Limiters............................................................................ 40 Flaps Load Relief ......................................................................... 40 Asymmetry of High Lift Devices ................................................... 40 Q-Feel .......................................................................................... 40 Control Divergence & Reversal.................................................... 42 Flutter & Control Surface Static Balance ..................................... 42 Buffeting....................................................................................... 44 Wing Structural Protection ........................................................... 44

Vortex devices & Vortex Lift................................................................ 46 Wing Fences ................................................................................ 46 Vortex Generators........................................................................ 46

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Vortilons ....................................................................................... 46 Saw-Tooth Leading Edges........................................................... 46 Super Critical Airfoils.................................................................... 46 Vortex Lift ..................................................................................... 48 Lift and Control From Strakes ...................................................... 48

Stall Warning....................................................................................... 50 Deep Stalling................................................................................ 52 Advanced Stall Warning Systems................................................ 52

The Canard Concept........................................................................... 54 Effects of Down-Springs & Bob Weights............................................. 56

Down-springs ............................................................................... 56 Bob Weights................................................................................. 56

11.1.1 Glossary of Terms & Abbreviations ......................................... 58

Edition: 1/MAR06 Revision: Responsible: NEP Module: TM11.1.1-2

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”THIS PAGE INTENTIONALLY LEFT BLANK”

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11.1.1 AEROPLANE AERODYNAMICS & FLIGHT CONTROLS The purpose with this submodule is to describe common practical ap-plications and the controlling of the aerodynamic forces introduced in submodule 8.3 and 8.4 (Basic Aerodynamics & Theory of Flight). This module does not explain from where the aerodynamic force origi-nates, but uses the laws of aerodynamics to explain the functions and purposes of; • Maneuvering flight controls, • Control principles, • Lift and drag inducing devices, • Boundary layer control (vortex devices), • Vortex lift, • Stall warning, • Aerodynamic protections, • Etc…

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RUDDER

ELEVATOR

AILERON

FLAPS

AILERONELEVATOR

AILERON

RUDDER

ELEVATOR

TRIMABLE HORIZONTALSTABILIZER (THS)

ROLL SPOILERS

GROUND SPOILERS(LIFT DUMPERS)

SPEED BRAKES

43

22

3

4

5

1

AILERON

SLATS

FLAPS

INTRODUCTION – LIGHT & HEAVY AIRCRAFT FLIGHT CONTROLS

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AIRCRAFT CONTROL Aircraft must be controlled in all three axes. For example, the pilot need to be able not only to set the pitch angle, but also to control the rate at which the angle changes. Aircraft motion is not limited only to motion about the main axes; an aircraft can also move in the plane along the axes (up/down, forward/rearward and transverse). Therefore, six degrees of freedom exist. There are several possible control systems which could be used to achieve this controllability. The Wright Brothers for instance obtained lateral control by "wing warping". They did this by twisting the wings to increase the angle of attack on one wing while reducing the angle of at-tack on the other wing, instead of conventional surface deflection. In generally the control system used depends on the whether the air-craft is: • Conventional, • Tailless (or delta), • Canard. A conventional (or classic) aircraft is one which has a separate main wing and horizontal tail and with the horizontal tail behind the main wing. An example of a conventional aircraft is shown below, with differ-ent tail configurations. Primary Flight Controls The most common control arrangement on the conventional aircraft is ailerons on the main wing and a horizontal tail known as the stabilizer with moveable elevators and a vertical fin (stabilizer) with a rudder. They are designed to change and control the moments about the roll, pitch, and yaw axes. The three different control surfaces are tradition-ally classified as the primary flight controls; • The elevator controls pitch or the longitudinal motion and thus is of-

ten called the longitudinal control,

• The ailerons control the roll or lateral motion and are therefore of-

ten called the lateral controls, • The rudder controls yaw or the directional motion and thus is called

the directional control. These control surfaces are flap-like surfaces that can be angularly de-flected at the command of the pilot. The figure hereunder introduces and defines the primary flight control surfaces. The ailerons are located at the trailing edge of the wing. Simi-larly, the elevator is located at the trailing edge of the horizontal stabi-lizer, and the rudder is at the trailing edge of the vertical stabilizer. The main control surfaces are attached to the airframe on hinges so they may move and thus deflect the air stream passing over them. This redirection of the air stream generates an unbalanced force to rotate the plane about the associated axis. Note: Other flight control surfaces can be added to the group of pri-mary flight controls if they have taken over a primary control function. Some aircraft achieves pitch control by a moveable horizontal stabilizer rather than an elevator. Other aircraft accommodate lateral control by roll spoilers (lateral control spoilers) rather than ailerons (or by combi-nation of the two). Secondary Flight Controls To the group of secondary flight controls we will traditionally find con-trol surfaces that are not directly involved with aircraft maneuvering, like; • Various tab controls (trim tabs, servo tabs, balance tabs, etc…), • Movable (trimable) stabilizers, • Spoilers/speed brakes, • Leading/trailing edge flaps, • Slats, etc…

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LONGITUDINAL/ROLL AXIS (X-AXIS)

CENTER OFGRAVITY (CG)

PITCH AXIS(Y-AXIS)YAW AXIS

(Z-AXIS)

REFERENCE AXIS

PITCHING

YAWING

ROLLING

AILERON (RH)

AILERON (LH)

ELEVATOR (LH)

ELEVATOR (RH)RUDDER

VERTICALSTABILIZER(FIN)

HORIZONTALSTABILIZER

CONVENTIONAL PRIMARY FLIGHT CONTROL LAYOUT

WING

CANARD(FIXED)

MAIN WING

TAIL PLANE (T-TAIL)

CANARD AIRCRAFT (PIAGGIO AVANTI)

DELTA AIRCRAFT (AEROSPATIALE/BAC CONCORDE)

DELTA AIRCRAFT WITH CANARD (EUROFIGHTER)

SLENDER DELTA(OGIVE WING)

CANARD(MOVABLE)

DELTAWING

CONVENTIONALTAIL DESIGN

T-TAILDESIGN

VERTICALSTABILIZER

HORIZONTALSTABILIZER

VENTRALFIN

DORSALFIN

TWIN TAIL

V-TAIL(BUTTERFLY TAIL)

AIRCRAFT CONTROLS

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CONTROL HANDLES AND LEVERS Traditionally aircraft are controlled by a column with a control wheel (also called a yoke) for elevator and aileron control. This can also be achieved by a control stick located in front of the seat or to the side. The rudder is controlled by the pedals. The mechanical (or electrical) connection between the controls and the flight control surfaces is not a topic for this submodule to describe. Pitch and Roll Controls Moving the column forward or rearward controls the pitch of the air-craft. Forward motion makes the nose drop; rearward motion makes the nose rise. The control wheel (attached to the column) controls the roll of the air-craft. Moving the wheel counterclockwise (CCW) makes the aircraft roll to the left, and a clockwise (CW) motion makes the aircraft roll to the right. The control stick works in the same manner. The stick is able to be cy-clic moved in a circumferential path in its pivoting attachment. Control sticks are typically used on military fighter aircrafts. The civilian Airbus A320 to 380’s uses a side stick instead of the traditional column. Yaw Control The two pedals are mechanical interconnected. When one pedal is pushed forward the other pedal is moved aft. A push on the left pedal makes the nose of the aircraft yaw to the left, and vise versa. The pedals also accommodate brake control by tipping the pedals for-ward with the tiptoes. Trim Controls Most aircraft can be trimmed around all three axes. The function of the trim control system in an aircraft is to ease the control force input re-quired of the pilot during steady flight, and reducing the workload. Trimming is typically achieved by turning or rotating a handle or by op-erating an electrical switch.

Smaller aircraft are typically trimmed by a mechanical operated handle. Larger aircraft are mostly trimmed by electrical actuation, from switches located on the control wheel or center pedestrial panel. The control stick sometimes comprises an electrical trim switch on the top of the grip, for pitch and roll trimming. Horizontal (Movable/Trimable) Stabilizer Control Pilot control input to a movable (trimable) horizontal stabilizer can be achieved either by input to a mechanical control wheel, located on the pedestrial panel (most large aircraft), or operating an electrical switch on the control wheel (yoke). Electrical (or digital) Flight Control Computer (FCC) systems are in many cases also provided with trim authorities of the stabilizer. Flaps and Slats Control Cockpit (flight deck) commands of the flaps and slats systems are mostly operated by handles or levers. Detents (or gated positions) al-low the pilot/flight crew accurately to set a specific position (in units or degrees). If slats are installed they will be operated in conjunction with the flaps handle/lever. Few aircraft facilitates independent control of flaps of slats (as a backup feature). Spoiler and Speed Brake Control Spoiler panels, installed on the upper side of the wings, are used as flight spoilers (roll and drag/speed-brake control) and ground spoilers (lift dumpers). Speed brakes are panels located on the wings or fuse-lage) that provides additional drag when deployed. A control handle/lever on the pedestrial panel controls the deflection of spoilers/speed brakes. If the panels also are used as ground spoilers an automatic deployment may occur at touch-down. If the flight spoilers take part in the roll control of the aircraft, a me-chanical or electrical signal is provided in proportion to input of the con-trol wheel (or stick).

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SIDESTICK

PEDALS

CONTROLWHEEL

CONTROLCOLUMN

RUDDERPEDALS

SPEED BRAKE/SPOILER LEVER

FLAP/SLATLEVER

HORIZONTALSTABILIZERTRIM WHEEL

STALL WARNINGSTICK SHAKER

CONTROLWHEEL

FLAPCONTROLLEVER

TYPICAL LIGHT AIRCRAFT CONTROLS (CESSNA)

SPEEDBRAKE/SPOILER LEVER

FLAPS/SLATSLEVER

HORIZONTALSTABILIZERTRIM WHEEL

MODERN FLY-BY-WIRE AIRCRAFTCONTROLS (AIRBUS)

B737NG AIRCRAFT CONTROLS

CONTROL WHEEL COMMANDS

NEUTRAL POSITION

LEFT ROLLRIGHT ROLL

CW

FWDPUSH

CCW

PITCHDOWN

PITCHUP

REARWARDPULL

RIGHTYAW

LEFTYAW

PEDAL COMMANDS

NEUTRALPOSITION

RUDDERPEDALS

YOKE CONTROL (BEAVER)

YOKE

CONTROL STICK (PILATUS)

TRIMWHEEL(S)

HORIZONTALSTABILIZERTRIM SWITCH

FLIGHT DECK/COCPIT CONTROLS & LEVERS

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THE FUNCTION OF FLIGHT CONTROL SURFACES Flight control surfaces that act as a deflectable part of the airfoil trailing edge are essential based on the same aerodynamic properties. The deflectable part, called a flap (not to be mistaken as a flaps sys-tem), regulates the effective shape (camber) of the airfoil and wing sec-tion. Aerodynamic Principles In the aerodynamic sense a change in camber will vary the effective angle of attack (AOA) and the local coefficient of lift, CL. A downward deflection of the control surface (aileron or elevator) causes an abrupt change in rearward curvature of the wing section. The low pressure region accelerates the airflow across the upper side of the wing, increasing the coefficient of lift (and the lift). A high pressure area is located on the lower side in the flap region. The overall result is an increase in the upward acting force. An upward deflection causes the opposite. The airflow is accelerated on the lower side and slowed down on the upper side. The result is an overall decrease in the upward acting force. A rudder which is given a sideward deflection will cause asymmetry in lift between the two sides of the vertical stabilizer (fin), creating a side-ward force. The change in force that is produced as the result of a certain deflec-tion is governed by the aerodynamic rules. Flying with low airspeed requires higher deflection angle to obtain a re-quired force; flying fast requires less. The density of air will also affect the force from the deflection. High alti-tude flying tends dampen the effect of a given deflection.

Maneuverability Motion is assumed to take place around the Center of Gravity (CG) which then involves forces and arms. Varying the force provided by horizontal stabilizer generates control moments about the pitch axis. Varying the force provided by the vertical stabilizer generates control moments around the yaw axis. Varying the differential left and right wing forces produces control moments about the longitudinal (roll) axis. The arm distances between the CG and the control forces highly af-fects the control moments. The CG will mostly change longitudinal but lateral changes may occur if because of fuel imbalance. Rearward CG location will reduce the yawing and pitching moment as-sociated with a given control surface deflection, because the effective control arm is reduced. Forward movement results in the opposite. Mechanical Stops Over-deflecting the control surface may cause boundary layer separa-tion making the rear portion of the airfoil stall. Mechanical stops pre-vent the pilot from applying too large deflections and loss of control. One set of mechanical stops (secondary stops) normally prevents too large control input, by limiting the input from control device (stick, yoke, column, wheel, pedals, etc.). Another set of stops (primary stop), lo-cated adjacent to the control surface, prevents too large deflections. Mechanical stops are adjustable allowing for rigging by the mainte-nance staff.

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MEAN CAMBER LINE

CHORD�

�

EFFECT OF CONTROL SURFACE (FLAP)

DOWNWARDDEFLECTION

�

NEUTRALAILERON

UPWARDDEFLECTION

INCREASEDLIFT

DECREASEDLIFT

NET LIFT (L)

NEUTRAL “FLAP”

NET LIFT (L)

�L

DOWNWARD DEFLECTION

NET LIFT (L)

�L

UPWARD DEFLECTION

CMAXL

0AOA

CL

0 10

0

1.0

1.6

EFFECT OF CHANGED CAMBER & AOA

RUDDERFORCE

ELEVATORFORCE

AILERONFORCE

CG

CONTROL FORCES AND ARMS

PITCHAXIS

ROLLAXIS

YAWAXIS

PRIMARY STOP(PRINCIPLE)

SECONDARY STOP(PRINCIPLE)

STICKMOVEMENT

DEFLECTIONRANGE

FUNCTION OF THE STOPS

THE AERODYNAMIC FUNCTION OF A FLIGHT CONTROL SURFACE

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Hinge Moments If an aerodynamic force acts on a control surface, it will try to rotate the control around its hinge in the direction of the force. Aerodynamic forces will seek to set the control in an equilibrium position. The hinge moment is a product of the force times the distance from the hinge line to the control surface center of pressure (CP). The force is due to the control surface area (S), its angular deflection and the dy-namic pressure (½ρv2). Note: The distance must be set perpendicular to the force. To move the control surface to a desired angular deflection and main-tain it in that position, the pilot has to overcome (balance) the hinge moment. This is done by applying a force to the cockpit control (stick force). The larger the control surface, deflection or dynamic pressure, the larger the stick force is required. Therefore the hinge moment gives the pilot a feel for the control. Direct Mechanical Control The simplest method of controlling the surface it the direct mechani-cal control. The pilot controls the surface by push-pull rods, pulleys and wires using only simple mechanical advantages (gear ratios), from levers, bell cranks, etc. An advantage of this control form is that the pilot has direct feel of the loadings on the aircraft, because the aerodynamic loadings is felt di-rectly in proportion with the hinge moments. This control form is typi-cally used on smaller aircraft operating with low airspeed and hinge moments not excessive to the pilot. For larger and faster aircraft the resulting aerodynamic forces could give hinge moments (stick forces) which would be too high for easy operation of the controls. In this cane the pilot will require assistance to move the controls in these conditions. This can be done either by using by using some form of aerodynamic balancing (achieved in a number ways) or powered (hydraulic) flight controls.

Aerodynamic Balancing Aerodynamic balancing involves using the aerodynamic forces on the control surfaces to reduce the hinge moments. a) An inset (or offset) hinge, meaning that the hinge line is moved

towards the aerodynamic force, will effectively reduce the hinge moment. An inset hinge does not reduce the effectiveness of the control.

b) Horn balance is another effective way to reduce the hinge moment.

By extending the control area forward of the hinge line, the force on this part will give hinge moments which are in opposite directions to the moments on main part of the surface. The overall hinge moment is therefore reduced without reducing the control effectiveness.

c) The balance panel (or internal balance) works similar to the inset

hinge, but the aerodynamic balance area is located inside the wing. Deflection of the control surface causes pressure changes on the

control, and these pressure changes are felt by the balance panel. Suppose that the control is deflected downwards to improve the

camber. The pressure above the airfoil is reduced and the pressure below increases. The pressure differential is felt across the balance panel. Recall that force equals pressure times area (F = P · A). The force produces a hinge moment which is in opposite directions to the hinge moment on main part of the surface. The overall hinge moment is therefore reduced. This type of balancing is sometimes used in conjunction with powered flight controls.

Tab Balancing Tab balancing is a specialized from of aerodynamic balancing, used on numerous aircraft of all categories. The control stick only moves a small part of the main control surface, thus, the stick forces remains low (even at high speed). The arm from the tab force to the hinge line produces the hinge moment that oper-ates the main control surface (it is an aerodynamic ratio changer).

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FORCE

HINGEMOMENT

ARM(HINGE

DISTANCE)

FORCE

REDUCEDHINGE MOMENT

REDUCEDARM

HINGE MOMENT

ANGULARDEFLECTION

INSET (OFFSET) HINGE

HINGELINE

HORNWING/STABILIZER

CONTROLSURFACE

HORN BALANCE

a

bF1

F2

HORNANGULARDEFLECTION

M2

M1

F1

a

F1

F2

M1

M1

M2

SERVO TAB PRINCIPLE

A)

B)

C)b

c

ELEVATOR UP COMMAND(SERVO TAB IS MOVED)

NEUTRAL (OR TRIMMED) CONDITION

F x b = F x c1 2

�M = 0

MAIN CONTROLSURFACE

TAB

CONTROLLINKAGE

BALANCE PANEL

FORCE

HIGH PRESSUREREGION

LOW PRESSUREREGION PRESSURE

DIFFERENTIAL

BALANCEPANEL

BALANCECHAMBER

ELEVATOR IS MOVEDBY SERVO TAB

THE SERVO TAB REDUCES THEREQUIRED STICK FORCE

HINGE MOMENTS & AERODYNAMIC BALANCING

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The tab is basically a small auxiliary control surface integral with the main control surface, and is used as a servo mechanism for the main surface, or as mean to avoid unwanted side effects from the deflection of the main control surface. Tabs are regarded as secondary flight controls (though it actually con-trol or biases a primary control surface). Tabs are classified into two main groups; • Control tabs, and • Geared (gear) tabs. Control tabs are found in three sub-groups; • Servo tabs, • Trim tabs (actually a fixed or trimable servo tab), and • Spring tabs. Geared tabs are found in two main groups; • Balance tabs, and • Anti-balance tabs. Note: An important detail is that control tabs are moved “before” the main surface is moved (the tab moves the main control), whereas a gear tab is moved by the main control (the gear tab is moved after). Servo tabs: Servo tabs are used to ease the required stick force ap-plied by the pilot. As explained the servo tab is an aerodynamic ratio changer. The control moments that the pilot applies (M1 = F1 x a) is proportion-ally less than the effective hinge moment applied to the main surface (M2 = F2 x c). The main surface will seek an equilibrium where F1 x b = F2 x c) which result in a certain deflection. Servo tabs are sometimes also seen on hydraulic powered flight con-trols, used for manual reversion if the normal hydraulic supply fails, providing sufficient servo assistance.

Trim tabs: The trim tab operates in the manner as the servo tab. The trim tab can be fixed or adjustable. The deflection of the trim tab produces a hinge moment that forces the main surface into a new neutral position, that otherwise was for the pi-lot to keep. Trimming is sometime accomplished by adjusting the servo tab control linkage. Spring tab: A disadvantage of the servo tab can be the lag of re-sponse when operating at low speed. The lowered dynamic pressure requires large control input to produce the required change in aerody-namic forces. In fact, direct control would be preferred at low speed. The spring tab combines the best from the two control methods. At low speed the spring forces exceeds the aerodynamic loads on the surface and the complete control operates, which is ideal for low speed operation. As the speed increase the aerodynamic loading on the con-trol increases, causing the spring pack to compress, thus allowing the servo tab to move. Geared tabs are either used to assist the deflection of the control sur-face or to counteract deflection, if, for some reasons, the control force is too excessive. A connecting rod is at one end attached to the tab and in the other end attached to the wing/stabilizer structure. A deflec-tion of the main surface will cause the geared tab to move caused by the geometry of the linkage. • The balance tab moves in the opposite direction of the main sur-

face, creating an aerodynamic force that aids the movement of the control (like a servo tab).

• The anti-balance tab moves in the same direction as the main sur-

face, creating a force that counteracts the movement of the control (increase the stick force). This can be used as a mean to dampen the defection of the main control (for instance used on many stabila-tor tail-planes). Or as a mean to increase the artificial feel (stick forces) as the deflection of a control surface increases.

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SPRING FORCE EXCEEDSAERODYNAMIC LOAD

LOW SPEED - LOW DYNAMIC PRESSURE(DIRECT CONTROL OF COMPLETE SURFACE)

HIGH SPEED - HIGH DYNAMIC PRESSURE(CONTROL OF SPRING TAB ONLY)

AERODYNAMIC LOADSPRING FORCE

EXCEEDS

- SERVO TAB

- TRIM TAB

- SPRING TAB

GEARED TAB

- BALANCE TAB

- ANTI BALANCE TAB

CONTROL TAB MOVES THE MAIN CONTROL SURFACE

IS BY THE MAIN SURFACEMOVED CONTROL

REDUCES THE STICK FORCE

REDUCES THE STICK FORCE AT HIGH SPEED

ADJUSTS THE NEUTRAL POSITION

REDUCES THE STICK FORCE

INCREASES THE STICK FORCE

SPRING TAB

SPRINGCONTROLINPUT

THE TAB FORCE MOVESTHE MAIN SURFACE

TABMAINSURFACE

CONTROL (OR GEAR) TAB

TRIM TAB

F1

F2

M1

M2

b

c

F x b = F x c1 2

�M = 0

SERVO TABF1

a

M1

TYPICAL TAB APPLICATION

MAIN CONTROLSURFACE

TAB

ADJUSTABLETRIM LINKAGE

TRIM KNOB/WHEEL(COCKPIT)

TRIM LINKAGEMD80 AILERON TABS

CONTROLTAB

TRIMTAB

TRIM TABS

FIXED RUDDER TRIM TAB

ANTI BALANCE TAB

BALANCE TAB

DECREASES THESTICK FORCES

INCREASES THESTICK FORCES

GEARED TABS TAB OVERVIEW TABLE

AERODYNAMIC BALANCING (CONTROL TABS & GEARED TABS)

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POWERED FLIGHT CONTROLS Powered flight controls are equivalent to hydraulic actuation. Hydraulic pressure at 3,000 PSI (and even in excess of that) is able to generate tremendous forces when exposed to a control area (piston area). Pow-ered flight control is classified into two sub categories; • Power assisted (power boosted), • Fully powered (fully in-reversible). Power Boosted Control The actuator assembly, - often called a Power Control Unit (PCU), has a main piston (or ram) hydraulic controlled by the small pilot oper-ated pilot/servo valve. The pilot valve controls the distribution of the supplied hydraulic pressure to the main piston. Power assisted/boosted control principle is similar to that used in many automobile vehicles. The cockpit control is mechanically connected to both the control surface (direct control) and the PCU pilot valve. An in-put command will move both the control surface and the pilot valve. Movement of the pilot valve makes the PCU assist moving the control linkage to the control. Note that the piston rod-end is attached to fixed structure. When hy-draulic pressure is applied to one side of the piston the complete PCU housing moves, also operating the control linkage to the surface. The housing also moves relative to the pilot valve and when the housing has moved the required distance (stroke) it has closed the pilot valve again. The relative motion between the housing and the pilot valve is called “follow-up” and allows for proportional control of the PCU. How much of the required stick force that comes from the pilot and how much that comes from the PCU depends on the boost ratio. A boost ratio of 5:1 means that the PCU delivers 80% of the required stick force, thereby reducing the required the cockpit. The mechanical geometry determines the boost ratio. The feedback from the control surface to the cockpit control is proportional to the boost ratio.

Power assisted/boosted systems often accommodate manual rever-sion which means that the pilot can attain full control of the control sur-face if the hydraulic supply fails (for instance by tab control). This re-quires that the two sides of the piston are interconnected and a mechanism that provide override of the input to the pilot valve. Fully Powered Control In a fully powered (in-reversible) system the cockpit control only oper-ates the servo valve in the PCU. In-reversible means that the pilot can deflect the control surface, but forces generated on the control surface can not be felt on the stick (or column/wheel/pedals). This is due to the hydraulic lock in the PCU. The PCU uses follow-up for proportional control. Artificial Feel & Centering Because the pilot only feels the load necessary to move the pilot valve in the PCU, an artificial feel system must be added to a powered con-trol the system. Artificial feel can simply be added by springs tension (or compression). The spring increases the stick forces when the de-flection increases giving the pilot a surrogate feeling of the aerody-namic load on the aircraft. The spring also provides centering of the servo mechanism so that the neutral position can be identified. Some controls have variable feel, achieved by a mechanical computer (Q-feel) that takes airspeed (and perhaps other data as well) into ac-count, to produce a more accurate “feel” for the pilot. Fly-By-Wire Principle Fly-By-Wire (FBW) control means that the mechanical control linkage to the PCU/actuator and the pilot valve has been replaced by an Elec-tro-Hydraulic Servo Valve (EHSV). The result is substantial weight savings. The EHSV controls the hydraulic fluid to the main piston when current excited from a driver in a digital (or analog) computer. The computer receives a feedback from an electric position sensor which provides proportional control of the PCU, by electrical follow-up.

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HYDRAULICSOURCE

P R

CONTROL INPUT FROM PILOT

CONTROL INPUT FROM PCU

PILOT VALVE

STRUCTURE

MAIN PISTON/RAM(STATIONARY)

COLUMN

PIVOT

LINKAGE

CONTROLSURFACE

POWER CONTROLUNIT (PCU)

HINGELINE

POWER BOOSTED (ASSISTED) CONTROL SYSTEM

FULLY POWERED (IRREVERSIBLE) CONTROL SYSTEM

HYDRAULICSOURCE

PR

TO PILOTVALVE

ARTIFICIAL FEEL & CENTERING MECHANISM

CAMROLLER SPRING

CAM SPRING ISCOMPRESSED

PROGESSIVEINCREASE INFEEL FORCE

OVALHOLE

NOTE: FEEDBACK FROM CONTROL

SURFACE TO COLUMN IN PROPOR-

TION TO THE BOOST RATIO

NOTE: NO FEEDBACK FROM CONTROL

SURFACE TO COLUMN BECAUSE OF A

HYDRAULIC “LOCK” IN THE PCU

HYDRAULICSOURCE

PR

FULLY POWERED FLY-BY-WIRE (FBW) CONTROL SYSTEM

SIDESTICK

STM

PCU FEED-BACK (LVDT)

EHSV (TM)

STEERING COMMANDS

STRUCTURE

PCU

ACTUATOR FEEDBACK (FOLLOW-UP)

SU

RFA

CE

PO

SIT

ION

FE

ED

BA

CKEXCITATION

Q-FEEL COMPUTER(FEEL INCREASESWITH AIRSPEED)

PITOT

STATIC

CAM

CONTROL LAW DATA(AIRDATA, AOA, WEIGHT,CG, A/C CONFIG, LOAD-FACTOR (n), ETC...)

AIRCRAFT MOTION

AUTOPILOT(AUTOFLIGHT)

POWERED FLIGHT CONTROLS (PRINCIPLES)

Aviation



LONGITUDINAL (PITCH) CONTROL Pitch control is a by far the most important control axis on the aircraft, because aircraft pitch, is closely related to control the angle of attack (AOA). The AOA is used to adjust for weight and CG changes, speed changes, different aircraft configurations (flaps, slat, etc.). Elevators Aircraft pitch control is traditionally achieved by the elevators. An ele-vator is mounted on the back edge of the horizontal stabilizer on each side of the fin. They move up and down together. Pulling back on the column, control wheel or stick will deflect the eleva-tors upward, making the aircraft pitch nose-up (ANU). This makes the wings fly at a higher angle of attack which generates more lift. Pushing the control forward will make the aircraft pitch nose down (AND). Cen-tering the control the elevators to neutral and stops the change of pitch. All Moving Tail Planes (Stabilizers) A stabilator is a combination of a stabilizer and an elevator. Instead of pure elevator motion the complete tailplane angle of incident can be changed. This increases or decreases the tailplane angle of attack; hence also the tail lift and aircraft pitch. The neutral point (angle of in-cidence) of the stabilator can be changed for trimming purposes. An anti-balance tab (a geared tab) is fitted to the Piper example be-low. The anti-balance tab produces an aerodynamic force that coun-teracts the direct pilot input to the stabilator. The tab makes sure that the “stick forces” gets proportionally higher and higher as the stabilator angle of incidence increases. It also pre-vents “aerodynamic run-away”. Recall that the location of the center of pressure changes with angle of attack (AOA). If the lift force of the sta-bilator attacks ahead of the pivot line; a moment assisting the input de-livered by the pilot could make the stabilator uncontrollable.

Supersonic military fighter aircraft are almost exclusively equipped with a pure all-moving tailplane which operates like the stabilator (without the tab), often referred as a slab tail. An important advantage of the all-moving tail is the large control mo-ments that can be created, even with small angle changes. An all-moving tail will in general produce less drag than a conventional stabi-lizer equipped with an elevator system. This has a particular interest when is comes to longitudinal trim forces. An all-moving tail performs far better in transonic and supersonic air-flows. Conventional elevator control will typically suffer control losses when operating in the transonic speed region, because of induced shock separation (see submodule 11.1.2 – High Speed). Shock separation near the trailing edge of a wing section will reduce the authority of any control surface located here (aileron, elevator, etc…). As the all-moving tail controls the lift over the entire tail area, it will not suffer as badly as an elevator operating in the wake.

Aviation



ELEVATOR

CG

AIRCRAFT PITCH CONTROL

PITCHAXIS NEUTRAL

PITCHINGMOTION ANU

AND

FIXED TAIL WITH ELEVATOR (CESSNA)

STABILATOR WITH BALANCE TAB (PIPER PA28)

HORNBALANCE

ELEVATOR

ANTI-BALANCETAB

STABILATOR

ALL-MOVING TAIL (F4 PHANTOM)

SLAB TAIL

ALL-MOVING TAIL

DEFLECTIONRANGE

TRANSONIC MACH NUMBER EFFECTS

SHOCK INDUCED SEPARA-TION CAUSES AERODYNAMICINEFFECTIVINESS

SHOCKWAVE

CONVENTIONAL CONTROL

SLAB TAIL CONTROL

SHOCK INDUCED SEPARATION

PITCH CONTROL

Aviation



High speed “tuck-under” (or mach tuck) is a phenomenon that occurs at high mach numbers because a shift in the location of the wing aero-dynamic center (AC) occurs. A supersonic lift distribution is somewhat more flattened out than the one of a subsonic. This results in a rear-ward motion of the center of pressure that, with respect to the CG, causes the aircraft to pitch down. The higher control authority of an all-moving tail makes it easier to counteract this by mach trimming the air-craft nose-up. If chock stalling of the tail occur an all-moving tailplane can easier be recovered, because the tailplane AOA effectively can be changed. All-Flying Tails Most jet transportation category aircraft have all-moving horizontal sta-bilizers, equipped also with a conventional elevator. This arrangement is often referred to as an all-flying tail. The elevator is used as a maneuvering control device coupled to the column/control wheel/stick (or autopilot) for pitch control. Trimming the aircraft for CG movement and aerodynamic phenomena’s (for instance mach trimming) is achieved by adjusting the stabilizer angle of inci-dence. The stabilizer is controlled by a screw-jack driven through a transmis-sion system. Manual control is possible by a cable-wire drive system. Automatic control is made possible either by electrical motors or by hy-draulic motors. The Flight Control Computer (FCC) system is often designed so that when a long term elevator deflection is sensed (by position sensors), the elevator deflection is cancelled out by auto-trimming of the horizon-tal stabilizer. A trimable horizontal stabilizer is a powerful control surface. A stabilizer run-away (uncommanded motion) caused by failure in the auto-flight systems, flight control computers (FCC’s) or stabilizer control circuit

could be fatal. In most aircraft it is therefore possible for the crew to manual override the electrical control circuit.

Aviation



ELEVATOR

TRIMABLESTABILIZER

BALANCE TAB

ALL-FLYING TAIL (B737)

ELEVATOR

PIVOT

STABILIZER

ALL-FLYING TAIL

TRIMRANGE

BEFORE TRIMMING

AFTER TRIMMING

TRIMABLE HORIZONTAL STABILIZER SYSTEM (TYPICAL)

M

COLUMNELECTRICAL TRIM SIGNALS

S

S

FCC’s

ELEVATORDEFLECTION

POS SIGNALS

STAB TRIMWHEEL

MANUAL TRIM SIGNALS

PITCH (ELEVATOR) COMMANDS

SCREWJACK

BALLNUT

STABDRIVE

FCC’s

MACHTRIM

HIGH SPEED TUCK-UNDER (MACH TRIMMING)

CGAC

- ML

LT

LOW MACH NUMBER TRIM

LOW SPEEDPRESSUREDISTRIBUTION

LT

CG AC

- ML

TRANSONIC MACH NUMBER TRIMHIGH SPEEDPRESSUREDISTRIBUTION

PITCH CONTROL

Aviation



DIRECTIONAL (YAW) CONTROL Yawing or directional control, accomplished by the rudder, is important in several aspects, such as; • Turn coordination, • Yaw damping (Dutch Roll), • Engine failure compensation (multi engine aircraft), • Runway alignment (de-crapping) during landing, etc… The rudder is mounted on the back edge of the fin in the tail. The rud-der works just like an elevator, except in a vertical plane. Deflection of the rudder to either side increases the camber of the fin and creates a force in opposite direction. This force, having an arm with respect to the CG, causes a yawing moment to act in the horizontal plane about the CG. When the pilot pushes the left pedal, the rudder deflects left. Pushing the right pedal causes the rudder to deflect right. Deflecting the rudder right pulls the tail left and causes the nose to yaw right and vise versa. Centering the rudder pedals returns the rudder to neutral and stops the yaw. Note: Asymmetrically engine thrust on multi-engine aircraft will also in-troduce yawing moments that must be outbalanced by a rudder input. Additionally, a necessary rudder deflection can be cancelled by asym-metrically engine thrust, in order to reduce the rudder trim drag. Turn Coordination Turning can be performed in a number of ways (however, most of them unpleasant for the crew and passengers). Turning must be made coor-dinated meaning that the force couple (gravity and the centrifugal force) acts along the vertical center line of the aircraft. Aircraft turn because of the banking created by the ailerons, not be-cause of a rudder input. Yawing is used to assist the banking to pre-vent side skidding (side slipping) during turning.

Turning will normally combine all three primary control systems; aileron to bank the aircraft, elevator to increase the AOA and keep up the nose during the turn and yaw pedals to make avoid coordinated turn. A coordinated turn is also called a balanced turn. Correct turning can most easy be understood by illustrating the turn and slip indicator. Correct turning is achieved when the “ball” in the slip indicator is lo-cated in the in the middle/bottom of the indicator. If the centrifugal force is dominant (to little pedal input) the ball moves outwards. If the too much pedal input is applied (the nose moves far too deep into the turn) the ball moves inwards. On larger aircraft turn coordination is mostly provided by the Flight Control Computers (FCC) and not the pilot. Yaw Damping Yaw damping is provided to dampen “Dutch Roll” (a dynamic instability mode) as explained in submodule 8.4 - Aircraft Stability & Dynamics. Dutch roll damping is carried out by a part of the FCC system or other flight augmentation system. Yaw damping is provided as a parallel input to the rudder actuation system as an electro/hydraulic command without any feedback to the pedals (which makes the pilot unaware of these rudder motions). Engine Failure Compensation Multi-engine aircraft, with lateral located engines (on the wings or at the tail) may use the rudder to outbalance the thrust asymmetry in case of engine failure. The size of the vertical stabilizer is typically related to the magnitude of the thrust loss and the additional drag associated with an engine failure. The distance between the AFT CG limit and the sta-bilizer is also an important design parameter. Multi engine propeller (turboprop) aircraft generates high drag if the propeller fails to feather in case of engine failure, which requires a large stabilizer and rudder area to outbalance.

Aviation



RUDDERFORCE

CG

AIRCRAFT YAW CONTROL

YAWAXIS

YAWING

RUDDERDEFLECTION

WINDCOMPONENT

RUNWAYHEADING

HEADING

WINDCOMPONENT

APPROACH WITHCROSS WIND

(RUDDER INPUT)

RUDDER

VERTICAL FIN/STABILIZER

RUDDERSIDE FORCE

ROLL(BANKS THEAIRCRAFT)

PITCH(KEEPS THEALTITUDE)

YAW(PREVENTSSKIDDING)

CONTROLING THE TURN (TURN COORDINATION)

CENTRIFUGALFORCE (SKID)

4 MIN TURN

TURN AND SLIPINDICATOR

NO SLIP/SKID

INSUFFICIENTYAWING (SKID)

TO MUCHYAWING

LH TURN(4 MIN)

W

FC

LW

WT

CENTRIFUGALFORCE

CENTRIPETALTURNINGFORCE

L =T FCL +W

CENTERLINE

TURNINDICATOR

SLIPINDICATOR

22

DE-CRAPPING(WITH RUDDER)

DIRECTIONAL (YAW) CONTROL

Aviation



LATERAL (ROLL) CONTROL Roll control can be achieved by several design methods. Traditionally ailerons are used to generate a rolling motion around the longitudinal axis. Another often used method is roll spoilers (lateral control spoil-ers) which are panels located on the upper wing surface, which can be deployed into the airflow lowering the wing lift. Aileron Roll Control Ailerons are mounted on the back edge of each wing near the wingtips, and move in opposite directions. The word aileron is French for "little wing". The ailerons change the effective airfoil camber of the part of the wing covered by the ailerons, and thereby the wing lift. With greater downward deflection, the lift will increase in the upward direction. An upward deflection causes the upward lift to decrease. A left or counterclockwise control input causes the lift on the right wing to increase and the lift on the left wing to decrease. When the pilot moves the stick left (or turns the control wheel counter-clockwise) the left aileron goes up and the right aileron goes down, and vise versa. This cause the aircraft to bank left and an unbalanced side force component of the wing lift force causes the aircraft's flight path to curve to the left. As previously explained aircraft turn because of the banking created by the ailerons, not because of a rudder input. Centering the stick returns the ailerons to neutral maintaining the bank angle. The plane will continue to turn until opposite aileron motion re-turns the bank angle to zero to fly straight. The actual character of the lateral stability will greatly influence on the behavior after the stick (or wheel) is centered. For both wings, the lift force of the wing section through the aileron is applied at the aerodynamic center of the section which is some dis-tance (an arm) from the aircraft center of gravity. This creates a mo-ment (torque) about the center of gravity. If the forces (and distances) are equal there is no net rolling moment on the aircraft. But if the forces are unequal, there is a net torque and the aircraft rotates about its cen-ter of gravity.

Roll Spoilers Some aircraft has roll control only accomplished by roll spoilers (no ai-lerons). The spoiler panel(s) will only deflect on the down-going wing causing the aircraft to roll by “lift dumping” (reducing the coefficient of lift – CL). A side effect is that the drag is increased only on the down-going wing, making it slow down compared to the up-going wing. At first glance this seems to be a disadvantage. But it is not, because it compensates for adverse yawing (see next text page). This arrangement makes movement of the spoiler dependent upon movement of the aileron. The linkage to the aileron is devised so that the spoiler is extended only when the aileron is raised. In other words, when the aileron moves downward, no deflection of the spoiler takes place. Aileron Reversal Large aileron deflections at high speed (high dynamic pressure) can be a serious treat for high speed aircraft. Wing sweep and taper causes the airfoil thickness to decrease towards the wing tips, reducing the tor-sional stiffness of the outboard wing section. A downward aileron mo-tion may actually reduce the angle of incidence (and the local AOA) so violently that the aircraft rolls to the opposite side than the com-manded. This phenomenon is called aileron reversal. An upward de-flection can cause the opposite effect. Flight testing must make sure that the aileron reversal speed is outside the normal speed envelope. One method to limit outboard wing torsion is by progressively lower the maximum aileron deflection angle as the speed increases, or by pre-venting aileron motion above a given speed. Lateral control may in-stead be achieved by another set of ailerons (called high speed ailer-ons or all-speed ailerons) located closer to the fuselage, where the wing structure is much more rigid. Frequently spoiler control is used as a mean to achieve high-speed lateral control. Both methods have the purpose to move the aerodynamic force inboard, to a position where the wing structure is more rigid.

Aviation



LATERAL CONTROL(ROLL)

AILERON (RH)

AILERON (LH)

WING

RH ROLL

RH ROLL

NEUTRAL POSITION

LEFT ROLL

RIGHT ROLL

“AFT LOOKING FORWARD (ALF)”

ROLL (LATERAL) SPOILERS(INITATES A RH ROLL)

PANAVIA TORNADO (PHOTO BY MICHAEL HIND)

HIGHSPEEDAILERON

LOWSPEEDAILERON

FLIGHT (ROLL)SPOILERS

ROLL/LATERAL CONTROL (BOEING B727)

TAILERONS

HIGH (ALL)SPEEDAILERON

LOWSPEEDAILERON

LARGEARMSMALL

ARM

FLIGHT (ROLL)SPOILERS

TORSION

AILERONLIFT

AOA ISLOWERED

AILERON REVERSAL

LATERAL (ROLL) CONTROL

Aviation



ADVERSE YAW An unwanted secondary effect of aileron operation is adverse yawing - a yawing moment in the opposite direction to the turn. In other words, using the ailerons to roll an aircraft to the right would produce a yawing motion to the left. Deflection of the ailerons introduces adverse yawing mainly because of two effects. The first effect arises from drag differences between the two wing halves caused by the changed wing geometry. This can be visualized this way; • The wing with the down moving aileron achieves increasing camber

increasing both lift and drag, and the opposite takes place on the wing with the up-moving aileron. The difference in lift makes the air-craft roll and the difference in drag make the aircraft yaw. Because the down-going wing has less drag it moves relatively faster than the up going wing making the yawing opposed to the roll. When moving the stick (or control wheel) to the left to bank the wings, ad-verse yaw moves the nose of the airplane to the right. This is called the secondary effect of the ailerons. Increasing the camber of a wing makes it increase profile drag and induced drag.

The second effect comes from the aerodynamic damping described in submodule 8.4 - Flight Stability and Dynamics; • A down-moving wing will experience an increase in AOA and vise

versa for an up-going. The down-moving wing increases its lift and the up going wing will have its lift decreased. As lift attack perpen-dicular to the resultant airflow the lift will act forward with respect to a vertical reference plane for the down going wing, which generates adverse yawing. For the up going wing the lift reduces acting behind a vertical reference plane with the same consequence. This is called the secondary effect of the roll.

Adverse yaw is more pronounced for light aircraft with long wings, such as gliders. It is counteracted by the pilot or the Flight Control Computer (FCC) system with the rudder.

Several design methods can be used by the manufacture to counter act adverse yaw, like; aileron differential control, rigging the ailerons with upfloat, frise ailerons, aileron-to-rudder interconnection or roll spoilers (lateral control spoilers). • Aileron differential control causes the upward moving aileron to a

higher deflection angle than the downward moving aileron. The higher deflection of the up-going aileron will make this wing produce additional drag whereas the lowered deflection of the down-going ai-leron lowers the drag, thus reducing the difference in drag between the two wings.

• Rigging the ailerons with up-float makes the aileron deflection be-

have as differential controlled ailerons, reducing the downward de-flection of the ailerons. This is used on many light aircraft.

• Frise ailerons have an asymmetric leading edge as shown in the

figure. The leading edge of the up-going aileron protrudes below the wing surface, causing higher drag. The leading edge of the down-going aileron remains shrouded and causes les drag.

• Flight spoilers (roll spoilers), only raised on the down-going wing,

will increase drag (and reduce lift) on this wing, thus reducing the difference in drag between the two wings. This arrangement makes movement of the spoiler dependent upon movement of the aileron. The linkage to the aileron is devised so that the spoiler is extended only when the aileron is raised. In other words, when the aileron moves downward, no deflection of the spoiler takes place.

• Aileron-to-rudder interconnection is a mechanical (or electrical)

input automatically provided to the rudder system, in proportion with the aileron command (which also provides turn coordination). This typically on many smaller aircraft.

Aviation



ROLL

ADVERSEYAW

RESTORINGROLL

LIFTDECREASES

LIFTINCREASES

ADVERSE YAW FROM ROLL (LATERAL DAMPING)

vF

vV

vRES

vRES

vF vVADVERSEYAW

ROLL

INCREASEDLIFT

INCREASEDDRAG

DECREASEDDRAG

DECREASEDLIFT

HIGHER INDUCED DRAG& PROFILER DRAG

ADVERSE YAW FROM AILERONS

AILERON DIFFERENTIAL

SMALL DOWN-WARD DEFLECTION

LARGE UPWARDDEFLECTION

FRISE AILERONS

HIGHER DRAG

RIGGING WITH “UP-FLOATING”

ROLL

DRAG

DRAG

DRAG

ROLL SPOILER

ROLL SPOILER COMPENSATION

EQUAL DRAG

ROLL/FLIGHT SPOILER

COUNTERACTING ADVERSE YAW

Aviation



LIFT INDUCING DEVICES The purpose with lift inducing devices (high lift device or lift augmen-tation) is to reduce the take-off and landing distance. This permits op-eration at greater weights from given runway (RWY) lengths and en-ables greater payloads to be carried. The term high lift device is somewhat misleading. Remember that lift must equal weight. The aircraft does not change weight just because a takeoff or a landing is to be performed. Instead the stall speed has to be reduced which requires that the maximum coefficient of lift (CLMAX) is increased. Slow speed device would therefore be a more relevant term. For a given weight the take-off and landing distances depends on the speed at the RWY “screen heights” specified by regulation (see sub-module 8.3 Aerodynamics). Regulation specifies safe margin to stall for at the screen heights (min 1.2VS and min 1.3VS) at takeoff and at land-ing). The stalling speed is determined by the CLMAX of the wing, and to obtain the lowest distances, the CLMAX must be as high as possible. Augmentation of CLMAX The lift coefficient is determined by the airfoil shape. Thickness (or fineness) ratio, leading edge radii and camber is important. Increasing the camber by leading and trailing edge devices increases the coeffi-cient of lift for a given angle of attack (AOA). It also increases CLMAX. For takeoff and landing a more cambered wing is desirable, but high drag will be the result at the cruising speed. The camber must there be variable so that the wing can be optimized for the flight condition. This is achieved by Trailing Edge (TE) flaps often combined with Leading Edge (LE) devices. Trailing Edge Flaps A flap is (as explained on page 10-11) is usually a hinged portion of the trailing edge on the airfoil section.

Flaps can however also be located at the leading edge (LE-flaps). The flap is usually used to positive increase the camber. However, jet fight-ers often use LE-flaps and TE-flaps for positive and negative variation of the camber. Lowering flaps increases both lift and drag, but not in the same propor-tion. Although the lift is the larger force, the proportional increase in drag is higher, and so the lift to drag ratio (L/D-ratio) decreases. The L/D-ratio is a measure of the aerodynamic efficiency and therefore flaps lowers the performance in areas such as range, endurance, climb angle and glide/decent angle. With flaps down the range is reduced (more fuel is used), the climb angle is reduced and the glide angle in-creased. In steady flight lift must equal the weight (L = W). Lowering the flaps in-creases the AOA for the same attitude. If the speed is kept constant the AOA must be reduced by lowered nose. Aircraft will therefore fly more nose-down with flaps extended, which is an advantage as it gives better visibility of the approach and landing. Trailing edge flaps changes the pressure distribution of the wing and therefore also the wing pitching moment, leading to trim changes when flaps are extended or retracted. Flaps also increase wing the down-wash, at the tail force produced by the tail plane. Increased camber tends to produce nose down pitching moments, increased downwash tends to do the opposite. The overall effect will be the sum of the two influenced by the type of flap, tail configuration, etc.

Aviation



L = ½ v C S�. . 2. .L

-2kgms

LIFT EQUATION

DENSITY:- ATMOSPHERIC PRESSURE,- TEMPERATURE (OAT)

WING AREA:- MAY BE VARIABLE BYFOWLER FLAPS

COEFFICIENT OF LIFT:- ANGLE OF ATTACK (AOA),- CAMBER (PROFILE SHAPE),- SURFACE CONDITION,- MACH NUMBER

AIR SPEED:- TRUE AIR SPEED (TAS)

V1

SCREENHEIGHTMIN 35’

VR

V2

50’

APPROACHPATTERN

LOW SPEED/HIGH DRAGCONFIG

MIN 1.3VS

TOUCH DOWN

HIGHTHRUSTSETTING

FINALAPPROACH

3 G/S0

MIN 1.2VS

“ONEENG

INE OUT (O

EO)”

TO PROF

ILE

“NORM

AL”

TOPRO

FILE

E = ½mvKIN2

LOW SPEED/LOW DRAGCONFIG

CLIMBSEGMENT

CLEAN (CRUISE) CONFIGURATION

TAKE-OFF CONFIGURATION

CHORD�

�

CAMBER VARIATION

APPROACH/LANDING CONFIGURATION

�

LOW SPEED/LOW DRAGCONFIG

LOW SPEED/HIGH DRAGCONFIG

LOW PRESSURE REGION

HIGH PRESSURE REGION

LOW SPEED CONFIG PRESSURE DISTRIBUTION

LIFT INDUCING DEVICES

Aviation



The basic principle of the trailing edge flap has been adapted in many ways. The more commonly used types are described hereunder. • The plain flap (b) is a simple construction and gives good increase

in CLMAX, although fairly high drag (low L/D-ratio). It is used mainly on low speed aircraft and where very short take-off and landing is not required.

• The split flap (c) forms a part of lower surface of the lower surface

of the trailing edge, the upper surface contour being unaffected when the flap is lowered. The split flap gives about the same in-crease in lift as the plain flap at low AOA, and slightly more at higher AOA as the upper surface camber in is not increased and so sepa-ration is delayed. The drag is however higher than for plain flaps because the wake is much stronger.

• Slotted flaps (d) have gaps (slots) allowing high pressure air from

the lower wing side to add energy to upper surface of the flap seg-ment. This adds energy to the tired boundary layer, postponing boundary layer separation to a higher AOA, increasing the CLMAX even further, as the deflection angle can be increased.

• Fowler flap (e) moves rearwards and then down when extended.

This gives an increase in camber and wing area. Typically fowler flaps are also slotted (single or multiple). Multiple slotted fowler flaps (f) has one or more vanes installed between the wing trailing edge and the flap section (providing the slot(s).

Multiple slotted fowler flaps have the best performance of all flap types, but are complex, heavy and volume requiring. Besides it takes time to extend/retract.

• The zap flap (g) is very similar to the fowler flap but without the slot. • Powered flaps generally use airflow from the engine to shape the

flow of air over the wing.

Blown flaps (h) use bleed air from the jet engine's compressor which is blown over the rear of the wing and flap, adding both air-flow and energy so that the flow remain attached at high AOA. In ef-fect the airflow acts as a sort of a slat for the flaps allowing higher flap deflection angle. Blown flaps are only used on few and mostly jet fighters. Thin winged fighters are not able to contain a complex and volume requiring slotted fowler flap system.

Another system uses the airflow from the engines directly, by plac-ing part of the flaps directly in the path of the exhaust, - a jet flap (i). Most aircraft have split the flap so the portion behind the engines does not move into the exhaust. But, if the flaps are made strong enough, the effects can be enormous. The C-17 Globemaster has a flap system using this principle (also called the Coanda effect).

Note: The action of the flaps is to increase the maximum coefficient of lift, CLMAX for a given AOA, but stalling AOA is reduced. Increasing the flap deflection angle increases both lift and drag. The comparison of the drag polar curves of the different types of flaps show a considerably difference in the lift and drag produced. Small de-flection angle will generally increase lift more than drag. At a given de-flection the increment in lift will be overtaken by larger and larger in-crement in drag. During takeoff, drag reduces the acceleration and climb performance and a small flap deflection is preferred to give as little drag as possible. For landing however, drag adds to the braking force and so flap drag is beneficial. Besides, the large flaps settings makes the aircraft fly more nose-down improving the vision during approach. High drag on approach requires high engine thrust. This is an advan-tage in case of a go-around of missed approach, because the crew does not have to wait long for the engine to accelerate to go-around thrust. As a part of the procedure drag is reduced by stepwise retrac-tion of the flaps as speed is gained, to improve the climb performance.

Aviation



PLAIN (SIMPLE) FLAP

SPLIT FLAP

BLOWN (BLC) FLAP

DOUBLE SLOTTED FOWLER FLAP

CLEAN/BASIC WING

A)

B)

C)

E)

D)

“DOUBLE SLOTTED”FOWLER FLAPSYSTEM

VANE

FLAP

H)

PLAIN

FOWLER

SLOTTED

SPLIT

BASIC

FOWLER

SLOTTED

SPLIT

PLAINBASIC

FOWLER

SLOTTED

SPLIT

PLAIN

CL

CD

AOA

CL

FLAPANGLE

�CL

ZAP FLAP

G)

F)

FOWLER FLAP

VERSUS ANGLE OF ATTACK

CMAXL

CMAXL

C ANDL

CD

C VERSUS (L/D-RATIO)L

C VERSUS DEFLECTIONL

SLOTTED FLAP

CESSNA (PLAIN) FLAP

FOWLER FLAP SYSTEM (BOEING B777)

SLOTTED FOWLER FLAP SYSTEM (BOEING B727)

JET FLAP (BY COANDA EFFECT)

SPLIT FLAP

SLOTS

SCREWJACKS

VANE

FLAPS

LEADING EDGELAP(LE) F

TRAILING EDGE(TE) FLAP

FLAP SYSTEM (A340)

TRAILING EDGE FLAPS

Aviation



Leading Edge Devices Lift inducing leading edge devices are groups into either; • Boundary Layer Controlling (BLC) devices, and • Camber varying devices. The group of Boundary Layer Controlling (BLC) devices primarily in-cludes fixed slots, automatic and variable slats. The group of camber varying (or controlling) devices includes Leading Edge flaps (LE-flaps), krueger flaps and variable camber LE-Flaps. Besides, the leading edge can be equipped with other features improv-ing the wing airflow when operating at low speed (like vortilons). All LE-device prevents the pressure distribution on the upper side from peaking at the leading edge. This delays separation to a higher AOA. LE-devices have only with very little effect on the wing pitching mo-ments. Slots and Slats The maximum coefficient of lift, CLMAX, can be increased through the use of leading edge slots. The low speed leading edge problem is due to the Reynolds number effects, explained in submodule 8.2 (Aerodynam-ics). The slot allows high pressure air from the lower surface to merge with upper airflow at the leading edge. The slot airflow causes a laminar boundary layer into a turbulent boundary layer, much more capable of overcoming wing curvature without flow separation. The airfoil can then be flown at a higher angle of attack before stall occurs and thus get a higher CLMAX value. For angles of attack less than the stall angle, however, the airfoil lift curve is relatively unaffected whether the slot is opened or closed. The slot can be a permanent (or fixed) channel in the wing leading-edge or formed by a leading-edge auxiliary airfoil called a slat. The slat is either fixed or variable. The opening of the slot can also be con-trolled by panels (rarely used method).

There are three types of slats – fixed, automatic or variable. With the fixed slat, the leading-edge slat is mounted a fixed distance from the airfoil. Its main disadvantage is that it creates excessive drag at high speeds. The automatic slat depends on the pressure distribution lifting the slat away from the wing at high angles of attack to open the slot. At low angles of attack (or high dynamic pressure), the slat is flush against the wing leading edge and reduces drag at high speeds com-pared with the fixed slot. Its main disadvantages are its added weight, complexity, and cost. Sometimes springs assist the aerodynamic forces in deploying the slat. Increasing the wing stall angle also has a disadvantage. To make use of the full CLMAX potential the aircraft must approach for a landing in an extreme nose-up attitude that reduces visibility. For that reason slat are seldom used alone, but is used in combination with trailing edge flaps. Note: The action of the slot/slat is very different than the flap, as it does not directly produce extra lift unless the AOA is increased. Its main pur-pose is to increase the maximum coefficient of lift, CLMAX. The variable slat is mechanically operated, typically sequenced in steps along with flaps extension or retraction. The variable slat is typi-cally a two-position slat (retracted - extended) or a three-position slat (retracted – extended – fully extended). Slats are mostly used on swept wings aircraft. Swept wings are used on aircraft flying with high mach numbers. Sweep increases the opera-tional mach number of the aircraft, because the local mach number of the wing is lowered. Slats are typically used on outer part (2/3) of the span, where the local coefficients of lift are highest (nature of a tapered swept wing).

Aviation



CMAXL

0AOA

CL

CLEA

NCO

NFIG

WIT

HFLA

PS

+SLA

TS

EFFECT OF SLAT AND FLAPS (TYPICAL)0

0

CMAXL

0AOA

CL

CLEA

NCO

NFIG

+ SLATS

0

0

EFFECT OF SLAT (TYPICAL)

PULLINGFORCE

PRESSUREDISTRIBUTION

NO OR TOO LATETRANSITION

EARLY FLOWSEPERATION(STALL)

LOW SPEED FLIGHT

LAMINARBOUNDARYLAYER

DISTRUBINGAIRFLOW

SLATSLOT TRANSITION TURBULENT

BOUNDARYLAYER

SLOT (SLAT) PRINCIPLE

SLOT

SLATSEGMENTS

VARIABLE SLAT

FIXED SLAT (OR SLOT)

MOVABLEPANEL

FIXED SLOT

McD DC8

VARIABLE SLAT (AIRBUS A340)

FIXEDSLOT

CRUISE

LOW SPEED

LOW AOA

HIGH AOA

FIXEDSLAT

KZ VII (FIXED SLOTS) FIESLER STORCH (FIXED SLATS)

RALLYE CLUB

AUTOMATIC SLAT

FLATPEAKY

SLATSCLEANCP CP

SLAT FULLYEXTENDED

SLAT FULLYRETRACTED

LEADING EDGE DEVICES - SLATS & SLOTS

Aviation



Leading Edge Flaps (Droop Nose) On high speed airfoils the leading edge may have very little camber and have a small nose radius. This can result in flow separation just aft of the leading edge even a fairly low AOA. The nature of a leading edge stall is much more violent than a stall that gradually spreads from the trailing edge, and must be avoided. Leading Edge (LE) flaps (or droop leading edges) increases the cam-ber of the airfoil nose section, causing the value of the maximum coef-ficient of lift, CLMAX, and the stall AOA to increase. LE-flaps are hinged section of the leading edge which can be deflected downwards. When the nose is drooped the airfoil meets the incoming airflow more direct, therefore preventing separation at the leading edge. The stall will be postponed to a higher AOA. Droop leading edges are often preferred on thin-winged aircraft (like fighter aircraft) not having sufficient space for the much more efficient slat system. There is no known use of hinged leading edges on com-mercial jet liners. Shaping the nose section with permanent (fixed) droop causes an ef-fect similar to LE-flaps. Permanent droop may however have a nega-tive impact on drag at higher speeds. Fixed droop is used on many smaller and medium sized aircraft. An airfoil with a droop nose is popu-larly referred to as a “droop snoot”. Krueger Flaps & Variable Camber LE-Flaps The krueger flap is a part of the lower surface of the leading edge, which can be rotated about its forward edge, as shown in the figure. When retracted the “extension” is stowed away in a cavity at the un-derside nose structure. To improve efficiency by adding a better leading edge profile, the cam-ber of the leading edge may be increases as it deploys, achieved by the variable camber leading edge flaps (also called a folding nose). To promote root stall (rather that tip stall) on a swept wing, krueger flaps are used on the inboard section because they are less efficient than variable camber leading edge flaps.

The variable camber leading edge flaps are therefore typically used in the outboard section (instead of slats). The krueger flap and variable camber leading edge flaps increases lift by increasing the airfoil camber (the coefficient of lift) and the wing a-rea. The Krueger flap is somewhat less effective than the slat but is probably simpler in mechanical design. Some aircraft employ slats on the outboard portion of the leading edge, where more powerful flow control is required, and krueger flaps on the inboard portion of the lead-ing edge. Krueger flaps are a typical Boeing feature. Unlike trailing edge flaps (and slats) which can be selected to interme-diate positions, the krueger flap is either stowed or deployed. Combining Slats & LE-Flaps Modern high speed transportation aircraft mostly uses a combination of TE-flaps and slats (eventually supported with krueger flaps in the root section). Configuring the wing for low speed must be made stepwise in order not to stall the wing when reducing the speed. On the other hand structural overspeed of the flaps can happen if extended at too high airspeed. Extending flaps will lower the wing stall AOA (which is progressively reduced as the flaps deflection increases). If a clean wing is operating with high AOA a flap deployment can in fact stall the wing! Leading edge devices will on the other hand increase the stall AOA. The lead-ing edge device must therefore be deployed before the trailing edge flaps are lowered. When retracting the flaps the leading edge devices must be the last part retracted. This is accomplished by automatic se-quencing. Some aircraft are however sequenced differently because of other aerodynamic characteristics and habits. Slats are often extended in two steps depending on the selected ex-tension. When extending a fowler flaps system the first part only con-cerns area reduction requiring partly slat extension. The downward de-flection comes later, requiring full slat extension.

Aviation



LEADING EDGE WITH FIXED DROOP(DROOP SNOOT)

LE-FLAPS

LEADING EDGE (LE) FLAPS (F16 FIGHTING FALCON)

LOW SPEED CONFIG

BI-CONVEX HIGH SPEED AIRFOIL (CLEAN)

DROOP LEADING EDGE

KRUEGER FLAP

RETRACTED

EXTENDED

VARIABLE CAMBERLEADING EDGE FLAP(FOLDING NODE)

BOEING B747

LEADING EDGE DEVICES ON BOEING B737

INBOARD KRUEGER FLAP (BOEING B737)

SLAT KRUEGER FLAP

SHOWN INTRANSIT....

RETRACTEDLE SLATSLE FLAPS(KRUEGER)

KRUEGERFLAP

VARIABLE CAMBERLE-FLAP

TE-FLAPS

FULL-EXTENDED

EXTENDED

BOEING 737 FLAPS/SLATS SEQUENCING

DOWN

FLAP

UPFLAP

40

30

25

15

10

5

2

1

0

FLAP

0AOA

CL

CLEA

NCO

NFIG

+LE

-FLA

PS

EFFECT OF LE-FLAPS (TYPICAL)0

0

CMAXL

FL 1 - 40

FL 10 - 40

FL 1 - 5

LEADING EDGE FLAPS (DROOP)

Aviation



DRAG INDUCING DEVICES Drag inducing devices, besides flaps, includes; • Spoilers, • Speed-brakes, • Drag chutes. The main difference between spoilers and speed-brakes is that spoil-ers are panels located on lift producing surfaces (the wings). Speed brakes are typically located on non-lifting parts (like the fuselage). Speed-brakes only provide drag. However, fighter aircraft that pro-duces fuselage lift will be affected both in terms of drag and lift when a large speed-brake panel is raised on the upper fuselage. Spoilers and speed brakes are typical “full aircraft envelope drag de-vices” which means that they are not structural speed limited, as for in-stance slats and flaps are. Spoilers Spoilers are used for decreasing wing lift; however, their specific de-sign, function, and use vary with different aircraft. The spoilers on some aircraft are long, narrow surfaces hinged at their leading edge to the upper wing skin. In the retracted position, the spoiler is flush with the wing skin. In the extended position, the spoiler is pivoted up and forward. The maximum deflection differs a lot but is seldom more than 50-60° above the hinge point. The spoilers disturb the smooth flow of air over the wing by causing boundary layer separation. The lift is consequently reduced, and considerable drag is added to the wing in proportion with the wake. Another type of spoiler (used by some older aircraft) is a long, slender, curved and perforated baffle that is raised edgewise through the upper surface of the wing forward of the aileron. It also disrupts the flow of air over the airfoil and destroys lift.

Spoilers are used as; • Roll spoilers (reducing lift and adding drag), to assist the ailerons

and preventing adverse yawing, • Speed brakes, primarily to adding drag, to lower the speed of the

aircraft or prevent the aircraft from building up excessive speed dur-ing decent (dive),

• Ground spoilers (reducing lift and adding drag) during the landing

run. Reducing lift increases the weight on the landing gear optimiz-ing braking; increasing drag assists the braking by unloading of the brakes and reducing the necessary landing run.

Note: The roll spoiler and speed brake spoiler function is typically achieved by the same part of the spoiler system. Dedicated spoiler panels, which are only used for the ground spoiler operation, are typi-cally prevented for deploying in flight by an interlock system. Speed-brakes Speed-brakes are hinged, movable control surfaces used for reducing the speed of aircraft, purely by adding drag. Some manufacturers refer to them as dive-brakes or dive-flaps. They are hinged to the sides, the top, bottom or rear of the fuselage. Regardless of their location, speed brakes serve the same purpose on all aircraft. Their primary purpose is to keep aircraft from building up excessive speed during decent/dive. They are also used in slowing down the speed of the aircraft prior to landing and during the landing run. Speed brakes are operated hydraulically or electrically. Drag Chutes Drag-chutes are almost exclusively used on high speed military aircraft to unload the landing gear and reduce the landing run. Its civil applica-tion is very, very limited.

Aviation



THE EFFECT OF GROUND SPOILERS

GROUNDSPOILERS

SPEEDBRAKES

SPEEDBRAKES

AVRO (BAC) 146

DRAG CHUTE (F117)

SPEED BRAKE (F16)

ROLL

DRAG

ROLL SPOILER(FLIGHT SPOILER)

DRAG

FLIGHT SPOILER(SPEED BRAKE)

DRAG

SPEED BRAKE (CESSNA A37 DRAGONFLY)

SPOILERPANELS

GROUND SPOILERS DEPLOYED (BOEING 737)

GROUNDSPOILERS

GROUNDSPOILER

FLIGHT/ROLL& GROUNDSPOILERS

SPOILER SYSTEM (BOEING 737NG)

FLIGHT SPOILERS DEPLOYED (AIRBUS A319)

INCREASED DRAGAND LIFT DUMP

SPOILER RETRACTED

THE EFFECT OF FLIGHT SPOILERS

SPOILER DEPLOYED

SPEED BRAKE (SUKHOI SU-27)

DEFLECTIONANGLE

SEPARATEDAIRFLOW

DRAG INDUCING DEVICES

Aviation



COMBINED FLIGHT CONTROLS Many aircraft are controlled by combined flight controls rather than the traditional controls previously introduced. Combined flight control sur-face is one that has more than just one control purpose (for instance to control pitch and roll or pitch and yaw). • V-tails (or butterfly tails) are used by aircraft to control pitch (eleva-

tor function) and yaw (rudder function). Examples are the older Beech Bonanza V35 and the modern F-117 Nighthawk. When a pure pitch command is given, both surfaces are deflected equally. A pure yaw command, say to the left, is achieved by de-flecting the left surface downward and the right upward (see figure). Combined yaw and pitch commands are carried out by differentially deflection.

• Flaperons combines flaps and ailerons. A single surface on each

wing serves both purposes: used as an aileron, the flaperons left and right are actuated differentially; when used as a flap, both flaperons are actuated downwards. Many fighter aircraft uses flaperons for roll control (F-16, F-18, etc…). Note that when a flaperon is actuated downwards (i.e. used as a flap) there is enough freedom of movement left to be able to still use the aileron function.

• Elevons combine elevators and ailerons and is used by delta

winged aircraft (like the Concorde or the JAS39 Grippen). When a pure pitch command is given, both elevons are deflected equally. A pure roll (lateral) command make the elevons deflect like aileron. Combined commands are carried out by differentially deflec-tion.

• Tailerons are used by many military fighter jets (like the F-14, F-16,

F-18, Tornado, etc.). Each of the two independent (variable inci-dence) horizontal stabilizers (slab tails), can be moved differentially (one up, one down, etc.).

Beside normal pitch control this also provides roll control. Tailerons are often used in combination with flaperons. The Tornado fighter for instance achieves roll control only by roll spoilers and tailerons. No traditional ailerons are installed. Modern canard equipped highly maneuverable fighters uses differ-entially controlled canards for multiple purposes, like pitch and roll control but also for aerodynamic braking during the landing roll-out.

• Spoiler panels may have several purposes. One is to serve as

speed brakes in flight, two to accommodate roll/lateral control (or secondary roll assisting the ailerons) and third to serve as ground spoilers (lift dumpers) during landing and rejected takeoff. The roll and speed brake function is often referred to as the flight spoiler function. Spoilers are often sectional div

Documents

This sub-module complies with Part-66 11.1.1 / CAT B1.1/1 ...€¦ · Aviation Module 11 Aerodynamics, Structures & Systems Module 11.1.1 Aeroplane Aerodynamics & Flight Controls/MAR06